Method for repairing a thermal barrier coating and repaired coating formed thereby

ABSTRACT

A method of repairing a thermal barrier coating on a component designed for use in a hostile thermal environment, such as turbine, combustor and augmentor components of a gas turbine engine. The method more particularly involves repairing a thermal barrier coating on a component that has suffered localized spallation of the thermal barrier coating. After cleaning the surface area of the component exposed by the localized spallation, a ceramic paste comprising a ceramic powder in a binder is applied to the surface area of the component. The binder is then reacted to yield a ceramic-containing repair coating that covers the surface area of the component and comprises the ceramic powder in a matrix of a material formed when the binder was reacted. The binder is preferably a ceramic precursor material that can be converted immediately to a ceramic or allowed to thermally decompose over time to form a ceramic, such that the repair coating has a ceramic matrix. The repair method can be performed while the component remains installed, e.g., in a gas turbine engine. Immediately after the reaction step, the gas turbine engine can resume operation during which the binder is further reacted/converted and the strength of the repair coating increases.

FIELD OF THE INVENTION

This invention relates to thermal barrier coatings for componentsexposed to high temperatures, such as the hostile thermal environment ofa gas turbine engine. More particularly, this invention is directed to amethod for repairing a thermal barrier coating that has sufferedlocalized spallation.

BACKGROUND OF THE INVENTION

Higher operating temperatures for gas turbine engines are continuouslysought in order to increase their efficiency. However, as operatingtemperatures increase, the high temperature durability of the componentsof the engine must correspondingly increase. Significant advances inhigh temperature capabilities have been achieved through the formulationof nickel and cobalt-base superalloys. Nonetheless, when used to formcomponents of the turbine, combustor and augmentor sections of a gasturbine engine, such alloys alone are often susceptible to damage byoxidation and hot corrosion attack and may not retain adequatemechanical properties. For this reason, these components are oftenprotected by an environmental and/or thermal-insulating coating, thelatter of which is termed a thermal barrier coating (TBC) system.Ceramic materials and particularly yttria-stabilized zirconia (YSZ) arewidely used as a thermal barrier coating (TBC), or topcoat, of TBCsystems used on gas turbine engine components. These particularmaterials are widely employed because they can be readily deposited byplasma spray, flame spray and vapor deposition techniques.

To be effective, TBC systems must have low thermal conductivity,strongly adhere to the component, and remain adherent throughout manyheating and cooling cycles. The latter requirement is particularlydemanding due to the different coefficients of thermal expansion betweenceramic topcoat materials and the superalloy substrates they protect. Topromote adhesion and extend the service life of a TBC system, anoxidation-resistant bond coat is often employed. Bond coats aretypically in the form of overlay coatings such as MCrAlX (where M isiron, cobalt and/or nickel, and X is yttrium or another rare earthelement), or diffusion aluminide coatings. During the deposition of theceramic TBC and subsequent exposures to high temperatures, such asduring engine operation, these bond coats form a tightly adherentalumina (Al₂O₃) layer or scale that adheres the TBC to the bond coat.

The service life of a TBC system is typically limited by a spallationevent brought on by thermal fatigue. Accordingly, a significantchallenge of TBC systems has been to obtain a more adherent ceramiclayer that is less susceptible to spalling when subjected to thermalcycling. Though significant advances have been made, there is theinevitable requirement to repair components whose thermal barriercoatings have spalled. Though spallation typically occurs in localizedregions or patches, the conventional repair method has been tocompletely remove the thermal barrier coating, restore or repair thebond layer surface as necessary, and then recoat the entire component.Prior art techniques for removing TBC's include grit blasting orchemically stripping with an alkaline solution at high temperatures andpressures. However, grit blasting is a slow, labor-intensive process anderodes the surface beneath the coating. With repetitive use, the gritblasting process eventually destroys the component. The use of analkaline solution to remove a thermal barrier coating is also less thanideal, since the process requires the use of an autoclave operating athigh temperatures and pressures. Consequently, conventional repairmethods are labor-intensive and expensive, and can be difficult toperform on components with complex geometries, such as airfoils andshrouds. As an alternative, U.S. Pat. No. 5,723,078 to Nagaraj et al.teach selectively repairing a spalled region of a TBC by texturing theexposed surface of the bond coat, and then depositing a ceramic materialon the textured surface by plasma spraying. While avoiding the necessityto strip the entire TBC from a component, the repair method taught byNagaraj et al. requires removal of the component in order to deposit theceramic material.

In the case of large power generation turbines, completely halting powergeneration for an extended period in order to remove components whoseTBC's have suffered only localized spallation is not economicallydesirable. As a result, components identified as having spalled TBC areoften analyzed to determine whether the spallation has occurred in ahigh stress area, and a judgment is then made as to the risk of damageto the turbine due to the reduced thermal protection of the component,which if excessive can lead to catastrophic failure of the component. Ifthe decision is to continue operation, the spalled component musttypically be scrapped at the end of operation because of the thermaldamage inflicted while operating the component without complete TBCcoverage.

Accordingly, it would be desirable if a repair method were availablethat could be performed on localized spalled areas of TBC on turbinehardware without necessitating that the component be removed from theturbine, so that downtime and scrappage are minimized.

BRIEF SUMMARY OF THE INVENTION

The present invention provides a method of repairing a thermal barriercoating on a component that has suffered localized spallation of thethermal barrier coating. After cleaning the surface area of thecomponent exposed by the localized spallation, a ceramic pastecomprising a ceramic powder in a binder is applied to the surface areaof the component. The binder is then reacted to yield aceramic-containing repair coating that covers the surface area of thecomponent and comprises the ceramic powder in a matrix of a materialformed when the binder was reacted. The binder is preferably a ceramicprecursor material that can be converted immediately to a ceramic orallowed to thermally decompose over time to form a ceramic, such thatthe repair coating has a ceramic matrix. According to the invention,each step of the repair method can be performed while the componentremains installed, e.g., in a gas turbine engine. Immediately after thereaction step, the gas turbine engine can resume operation to furtherreact/convert the binder, by which the strength of the repair coatingincreases.

In view of the above, it can be appreciated that the invention overcomesseveral disadvantages of prior methods used to repair thermal barriercoatings. In particular, the method of this invention does not requirethe thermal barrier coating to be completely removed, nor does theinvention require removal of the component in order to repair itsthermal barrier coating. As a further advantage, the repair process doesnot require a high temperature treatment, since the repair coatingexhibits sufficient strength to withstand engine operation, during whichtime the precursor binder is gradually converted to form a ceramicmatrix. As a result, minimal downtime is necessary to complete therepair and resume operation of the turbine engine. In the case of largepower generation turbines, the cost is avoided of completely haltingpower generation for an extended period in order to remove, repair andthen reinstall a component that has suffered only localized spallation.Also avoided is the need to decide whether or not to continue operationof the turbine until the spalled component is no longer salvageable atthe risk of damaging the turbine.

Other objects and advantages of this invention will be betterappreciated from the following detailed description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional representation of a component surfaceprotected by a thermal barrier coating that has suffered localizedspallation.

FIGS. 2 and 3 are cross-sectional representations of the componentsurface of FIG. 1 during the repair of the thermal barrier coating inaccordance with the present invention.

DETAILED DESCRIPTION OF THE INVENTION

The present invention is directed to components protected by thermalbarrier coatings for operation within environments characterized byrelatively high temperatures, and therefore subjected to severe thermalstresses and thermal cycling. Notable examples of such componentsinclude the high and low pressure turbine nozzles and blades, shrouds,combustor liners and augmentor hardware of gas turbine engines for usein aircraft and industrial applications. While the advantages of thisinvention are particularly applicable to components of gas turbineengines, the invention is generally applicable to any component in whicha thermal barrier coating is used to thermally insulate a component fromits environment.

Represented in FIG. 1 is a surface region of a component 10 protected bya thermal barrier coating (TBC) system 12. The TBC system 12 is shown asbeing composed of a bond coat 14 formed on the surface of the component10, and a ceramic layer 16 deposited on the bond coat 14 as the thermalbarrier coating. As is the situation with high temperature components ofgas turbine engines, the component 10 may be formed of a nickel, cobaltor iron-base superalloy. The bond coat 14 is preferably formed of ametallic oxidation-resistant material, so as to protect the underlyingcomponent 10 from oxidation and enable the ceramic layer 16 to moretenaciously adhere to the component 10. Suitable materials for the bondcoat 14 include MCrAlX overlay coatings and diffusion aluminidecoatings. Following deposition of the bond coat 14, an oxide scale 18forms on the surface of the bond coat 14 at elevated temperatures. Theoxide scale 18 provides a surface to which the ceramic layer 16 moretenaciously adheres, thereby promoting the spallation resistance of theceramic layer 16.

The ceramic layer 16 is represented as having been deposited by plasmaspraying, such as air plasma spraying (APS). A preferred material forthe ceramic layer 16 is an yttria-stabilized zirconia (YSZ), a preferredcomposition being about 4 to about 8 weight percent yttria, though otherceramic materials could be used, such as yttria, nonstabilized zirconia,or zirconia stabilized by magnesia (MgO), ceria (CeO₂), scandia (Sc₂O₃)and/or other oxides. The ceramic layer 16 is deposited to a thicknessthat is sufficient to provide the required thermal protection for thecomponent 10, typically on the order of about 50 to about 300micrometers for most gas turbine engine components.

As a gas turbine engine component, surfaces of the component 10 aresubjected to hot combustion gases during operation of the engine, andare therefore subjected to severe attack by oxidation, corrosion anderosion. Accordingly, the component 10 must remain protected from itshostile operating environment by the TBC system 12. Loss of the ceramiclayer 16 due to spallation leads to premature and often rapiddeterioration of the component 10. A localized spalled region 20 of theceramic layer 16 is represented in FIG. 1, with the TBC repair processof this invention being represented in FIGS. 2 and 3. According to theinvention, each of the following steps performed in the repair of thecomponent 10 is performed while the component 10 remains installed inthe turbine engine, thereby completely avoiding the prior requirement toremove and later reinstall the component.

The repair process begins with cleaning the surface 22 exposed by thelocalized spalled region 20 so as to remove loose oxides andcontaminants such as grease, oils and soot, though preferably withoutdamaging the oxide scale 18 or removing any residual fragments of theceramic layer 16 that adhere to the scale 18. While various techniquescan be used, a preferred method is to remove loose materials for thespalled region 20, and then clean the surface 22 with alcohol and/oracetone. This step can be selectively performed to ensure that thesurrounding undamaged ceramic layer 16 is not subjected to theprocedure.

Once free of contaminants, the spalled region 20 is filled with aceramic paste 24, as represented by FIG. 2. According to the invention,the ceramic paste 24 is a mixture of ceramic powders and a binder thatwhen sufficiently heated forms a ceramic repair coating 26 shown in FIG.3 as adhering to the surface 22, which may be defined by portions of thebond coat 14, oxide scale 18 and/or remnants of the ceramic layer 16.The ceramic powder preferably is a mixture of one or more refractoryoxides, with preferred oxides including alumina, zirconia, hafnia,magnesia and silica. The binder is a ceramic precursor material,preferably a silicone or a phosphate-based composition, though it isforeseeable that other ceramic precursor binders could be used,including sol gel chemistries that thermally decompose to formrefractory oxides, and possibly calcium aluminate cements. According toa preferred embodiment of the invention, the ceramic powder containsabout 5 to about 85 weight percent alumina, 0 to about 60 weight percentzirconia, 0 to about 40 weight percent silica, 0 to about 55 weightpercent hafnia, 0 to about 55 weight percent magnesia and 0 to about 25weight percent zinc titanate. A ceramic powder that has been foundparticularly suitable contains about 42 weight percent alumina and about58 weight percent zirconia. The ceramic powder is combined with thebinder and a solvent in an amount sufficient to constitute about 50 toabout 95 weight percent of the resulting ceramic paste 24. A powder tobinder ratio of about 3 to 1 is generally preferred, such as when usingthe above-noted ceramic powder containing about 42 weight percentalumina and about 58 weight percent zirconia.

Preferred silicone binders include resins manufactured by GE Siliconesunder the names SR350 and SR355, and classified as amethylsesquisiloxane mixture of the polysiloxane family. These bindersare preferably used in amounts of up to about 45 weight percent of theceramic paste 24. Preferred phosphate-based binders include aluminumphosphate and complex phosphate materials that are commerciallyavailable from various sources such as Budenheim, Chemische Fabrik, inamounts of up to about 20 weight percent of the ceramic paste 24. Thesolvent content of the paste 24 will depend on the particular binderused, with the amount being sufficient to dissolve the binder. Asuitable solvent for the preferred silicone and phosphate-based bindersis an alcohol such as denatured alcohol (e.g., ethyl alcohol combinedwith 5% isopropyl alcohol) in an amount of about 5 to 65 weight percentof the paste 24.

The ceramic paste 24 may include additional additives, particularly oneor more surfactants to achieve a suitably tacky consistency that enablesthe paste 24 to adhere to the composition at the surface 22, which asnoted above may be defined by portions of the metallic bond coat 14, theoxide scale 18 and/or remnants of the ceramic layer 16. For example, upto about 10 weight percent of a nonionic surfactant may be desirable.Examples of suitable surfactants commercially available are P521A andMerpol from Witco and Stephan, respectively.

The paste 24 can be applied in any suitable manner, such as with atrowel. Depending on its composition, the binder of the paste 24 mayreact at room temperature, or its reaction accelerated by heating suchas with a heat lamp, torch or other heat source until the strength ofthe resulting repair coating 26 has reached a required level foroperation in the turbine engine. Suitable thermal treatments are aboutsixteen hours at room temperature to cure a silicone binder, and abouttwo hours at about 150° C. to react a phosphate-based binder. Duringoperation of the turbine engine, the repair coating 26 continues toreact, associated with an increase in the strength and other mechanicalproperties of the coating 26. In the case where a silicone binder isused, the binder initially cures by polymerization to form a siliconematrix whose strength is sufficient for engine operation. With extendeduse at high temperatures, the silicone thermally decomposes to silica,forming a silica matrix in which the particles of the ceramic powder aredispersed. In the case where a phosphate-based binder is used, thebinder chemically modifies the surfaces of the ceramic powder particlesduring the thermal treatment to form complex phosphate glasses andphosphate bonds. The complex glasses bind the powders together and forma matrix of high melting point phosphate glasses. With further exposureto high temperatures, such as during engine operation, thephosphate-based binder eventually migrates into the powder particles toform ceramic-to-ceramic bonds. Testing of repair coatings 26 withsilicone as the binder has shown that, through the process of beingconverted from an unfired cured polymer matrix composition to a fullyfired ceramic composition, the repair coatings 26 of this invention arecharacterized by enough residual strength to remain firmly adhered tothe surface 22 within the spalled region 20 in the ceramic layer 16.

While the invention has been described in terms of a preferredembodiment, it is apparent that other forms could be adopted by oneskilled in the art. Accordingly, the scope of the invention is to belimited only by the following claims.

What is claimed is:
 1. A method for repairing a thermal barrier coatingon a component that has suffered localized spallation of the thermalbarrier coating, the method comprising the steps of: applying a ceramicpaste on a surface area of the component exposed by the localizedspallation, the ceramic paste comprising a ceramic powder in a binder,the ceramic powder comprising alumina and zirconia, the binder beingchosen from the group consisting of ceramic precursor binders thatthermally decompose to form a refractory material; and then heating thebinder to yield a repair coating that covers the surface area of thecomponent, the repair coating comprising the ceramic powder in a matrixthat comprises the refractory material formed by reacting the binder. 2.A method according to claim 1, further comprising the step of cleaningthe surface area of the component prior to the applying step so as toremove contaminants without removing any adherent residual fragments ofthe thermal barrier coating.
 3. A method according to claim 1, whereinthe ceramic powder further comprises at least one ceramic materialchosen from the group consisting of hafnia, magnesia and silica.
 4. Amethod according to claim 1, wherein the ceramic powder consists ofabout 5 to about 85 weight percent alumina, 0 to about 40 weight percentsilica, up to about 60 weight percent zirconia, 0 to about 55 weightpercent hafnia, 0 to about 55 weight percent magnesia and 0 to about 25weight percent zinc titanate.
 5. A method according to claim 1, whereinthe ceramic powder comprises about 42 weight percent alumina and about58 weight percent zirconia.
 6. A method according to claim 1, whereinthe binder is a silicone and constitutes up to about 45 weight percentof the ceramic paste.
 7. A method according to claim 1, wherein thebinder is a phosphate-based composition and constitutes up to about 20weight percent of the ceramic paste.
 8. A method according to claim 1,wherein the component is installed in a gas turbine engine.
 9. A methodaccording to claim 8, wherein the method is performed while thecomponent remains installed in the gas turbine engine.
 10. A methodaccording to claim 1, wherein the gas turbine engine is operated afterthe reaction step during which the binder further reacts and the matrixconsists of a ceramic material.
 11. A method according to claim 1,wherein the thermal barrier coating is yttria-stabilized zirconiadeposited on a metallic bond coat on the component by plasma spraying.12. A method for repairing a thermal barrier coating on a componentinstalled in a gas turbine engine and which has suffered localizedspallation of the thermal barrier coating so as to expose a surface areadefined at least in part by an oxide scale on the component, the methodcomprising the steps of: without removing the component from the gasturbine engine, cleaning the surface area of the component exposed bythe localized spallation so as to remove contaminants without damagingthe oxide scale; applying a ceramic paste on the surface area of thecomponent, the ceramic paste comprising a ceramic powder in a ceramicprecursor binder dissolved in a solvent, the ceramic powder comprisingat last one ceramic material chosen from the group consisting ofalumina, zirconia, hafnia, magnesia and silica, the binder being chosenfrom the group consisting of phosphate-based compositions; evaporatingthe solvent and heating the ceramic paste to yield a ceramic-containingrepair coating that covers the surface area of the component, the repaircoating comprising the ceramic powder in a matrix consisting of aphosphate glass material formed by thermal decomposition of the binder;and then operating the gas turbine engine during which the phosphateglass material forms ceramic-to-ceramic bonds with the ceramic powder.13. A method according to claim 12, herein the ceramic powder consistsof about 5 to about 85 weight percent alumina, 0 to about 40 weightpercent silica, 0 to about 60 weight percent zirconia, 0 to about 55weight percent hafnia, 0 to about 55 weight percent magnesia and 0 toabout 25 weight percent zinc titanate.
 14. A method according to claim12, wherein the ceramic powder consists of about 42 weight percentalumina and about 58 weight percent zirconia.
 15. A method according toclaim 12, wherein the ceramic powder comprises alumina and zirconia. 16.A method according to claim 12, wherein the binder constitutes up toabout 20 weight percent of the ceramic paste.
 17. A method according toclaim 13, wherein the thermal barrier coating is yttria-stabilizedzirconia deposited on a metallic bond coat on the component by plasmaspraying.